Vane assembly of a gas turbine engine

ABSTRACT

A first stage vane array of a high pressure turbine that may be for a geared turbofan engine includes a plurality of airfoils circumferentially spaced from one-another and orientated about an engine axis. Each airfoil has a leading edge and a trailing edge with the trailing edge being circumferentially separated by the next adjacent trailing edge by a pitch distance. The leading a trailing edges of each one of the plurality of airfoils are axially separated by an axial chord length. A pitch-to-chord ratio of the pitch distance over the axial chord length is equal to or greater than 1.7.

This application is a continuation of U.S. patent application Ser. No.14/938,484 filed Nov. 11, 2015, which claims priority to U.S. PatentAppln. No. 62/089,014 filed Dec. 8, 2014, which are hereby incorporatedby reference.

BACKGROUND

The present disclosure relates to a gas turbine engine, and moreparticularly, to a first stage vane assembly of a high pressure turbineof the engine.

Gas turbine engines are rotary-type combustion turbine engines builtaround a power core made up of a compressor, combustor and turbine,arranged in flow series with an upstream inlet and downstream exhaust.The compressor section compresses air from the inlet, which is mixedwith fuel in the combustor and ignited to generate hot combustion gas.The turbine section extracts energy from the expanding combustion gas,and drives the compressor section via a common shaft. Expandedcombustion products are exhausted downstream, and energy is delivered inthe form of rotational energy in the shaft, reactive thrust from theexhaust, or both.

The turbine section typically includes alternating rows of turbine vanesand turbine blades. The turbine vanes are stationary and function todirect the hot combustion gases that exit the combustor section. Thevanes and blades each project from respective platforms that whenassembled form vane and blade rings. Airfoils of the vane and bladerings are designed with pitch-to-chord ratios that are generallydependent on a wide variety of engine characteristics and operatingparameters. Achieving the optimal pitch-to-chord ratio is desirable tooptimize engine efficiency and performance. Moreover, achieving higherpitch-to-chord ratios may reduce the number of required airfoils and/orreduce cooling requirements thereby improving engine efficiency andreducing engine manufacturing and maintenance costs.

SUMMARY

A first stage vane assembly of a high pressure turbine of a gas turbineengine according to one, non-limiting, embodiment of the presentdisclosure includes a first airfoil configured to be circumferentiallyspaced from an adjacent second airfoil and orientated about an engineaxis, wherein the first airfoil has a leading edge and a trailing edgewith the trailing edge configured to be circumferentially separated fromthe trailing edge of the second airfoil by a pitch distance, and theleading and trailing edges of the first airfoil are axially separated byan axial chord length, and wherein a pitch-to-chord ratio of pitchdistance over axial chord length is equal to or greater than 1.7.

Additionally to the foregoing embodiment, the first airfoil has athickness-to-axial chord ratio that is greater than forty percent.

In the alternative or additionally thereto, in the foregoing embodiment,the thickness-to-axial chord ratio is about fifty-three percent.

In the alternative or additionally thereto, in the foregoing embodiment,the trailing edge has an angle that is greater than seventy-fivedegrees.

In the alternative or additionally thereto, in the foregoing embodiment,the pitch-to-chord ratio is within a range of about 1.7 to 2.0.

In the alternative or additionally thereto, in the foregoing embodiment,the pitch-to-chord ratio is about 1.8.

In the alternative or additionally thereto, in the foregoing embodiment,the assembly includes an inner endwall, wherein the first airfoilprojects radially outward from an outward surface of the inner endwalland the outward surface includes at least in-part a concave regionlocated axially between the leading and trailing edges andcircumferentially adjacent to the first airfoil.

In the alternative or additionally thereto, in the foregoing embodiment,the outward surface includes a convex region proximate to the leadingedge of the first airfoil.

In the alternative or additionally thereto, in the foregoing embodiment,the assembly includes an outer endwall, wherein the first airfoilprojects radially inward from an inward surface of the outer endwall andthe inward surface includes at least in-part a concave region locatedaxially between the leading and trailing edges and circumferentiallyadjacent to the first airfoil.

In the alternative or additionally thereto, in the foregoing embodiment,the inward surface includes a convex region proximate to the leadingedge of the first airfoil.

A first stage of a high pressure turbine of a gas turbine engineaccording to another, non-limiting, embodiment includes a vane arrayhaving a plurality of airfoils circumferentially spaced from one-anotherand orientated about an engine axis, wherein each one of the pluralityof airfoils have a leading edge and a trailing edge with each one of thetrailing edges being circumferentially separated by the next adjacenttrailing edge by a pitch distance, and the leading and trailing edges ofeach one of the plurality of airfoils are axially separated by an axialchord length, and wherein a pitch-to-chord ratio of pitch distance overaxial chord length is equal to or greater than 1.7.

A turbofan engine according to another, non-limiting, embodimentincludes a fan configured for rotation about an engine axis; a lowpressure compressor; a low pressure turbine; a low spool mounted forrotation about the engine axis and interconnecting the low pressurecompressor and the low pressure turbine; a high pressure compressor; ahigh pressure turbine including a first stage vane array having aplurality of airfoils spaced circumferentially from one-another andhaving a pitch-to-chord ratio greater than 1.7; and a high spool mountedfor rotation about the engine axis and interconnecting the high pressurecompressor and the high pressure turbine.

Additionally to the foregoing embodiment, the pitch-to-chord ratio iswithin a range of about 1.7 to 2.0.

In the alternative or additionally thereto, in the foregoing embodiment,the turbofan engine includes a geared architecture configured betweenthe fan and the low spool.

In the alternative or additionally thereto, in the foregoing embodiment,the first stage vane array includes an annular inner endwall having anoutward surface generally facing radially outward and having a pluralityof concave regions with each one of the plurality of concave regionslocated between respective adjacent airfoils of the plurality ofairfoils that extend radially outward from the outward surface.

In the alternative or additionally thereto, in the foregoing embodiment,the outward surface includes a plurality of convex regions with each oneof the plurality of convex regions proximate to a respective leadingedge of the plurality of airfoils.

In the alternative or additionally thereto, in the foregoing embodiment,the first stage vane array includes an annular outer endwall having aninward surface generally facing radially inward and having a pluralityof concave regions with each one of the plurality of concave regionslocated between respective adjacent airfoils of the plurality ofairfoils that extend radially inward from the inward surface.

In the alternative or additionally thereto, in the foregoing embodiment,the inward surface includes a plurality of convex regions with each oneof the plurality of convex regions proximate to a respective leadingedge of the plurality of airfoils.

In the alternative or additionally thereto, in the foregoing embodiment,each airfoil of the plurality of airfoils have a thickness-to-axialchord ratio that is greater than forty percent.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand figures are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiments. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross section of an exemplary gas turbine engine;

FIG. 2 is an exploded perspective view of a vane array of a first stageof a high pressure turbine of the gas turbine engine;

FIG. 3 is a schematic of adjacent airfoils of the vane arrayillustrating pitch and chord length relationships;

FIG. 4 is a partial perspective view of the vane array; and

FIG. 5 is a partial plan view of an inner endwall of the vane array.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20 disclosed as atwo-spool, geared, turbofan engine that generally incorporates a fansection 22, a compressor section 24, a combustor section 26 and aturbine section 28. Alternative engines might include an augmentorsection (not shown) among other systems or features. The fan section 22drives air along a bypass flowpath while the compressor section 24drives air along a core flowpath for compression and communication intothe combustor section 26, then expansion through the turbine section 28.Although depicted as a turbofan in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with turbofans as the teachings may be applied toother types of turbine engine architecture such as turbojets,turboshafts, and three-spool (plus fan) turbofans with an intermediatespool.

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine static structure 36 or engine case via severalbearing structures 38. The low spool 30 generally includes an innershaft 40 that interconnects a fan 42 of the fan section 22, a lowpressure compressor 44 (“LPC”) of the compressor section 24 and a lowpressure turbine 46 (“LPT”) of the turbine section 28. The inner shaft40 drives the fan 42 directly or through a geared architecture 48 todrive the fan 42 at a lower speed than the low spool 30. An exemplaryreduction transmission is an epicyclic transmission, namely a planetaryor star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor 52 (“HPC”) of the compressor section 24 and highpressure turbine 54 (“HPT”) of the turbine section 28. A combustor 56 ofthe combustor section 26 is arranged between the HPC 52 and the HPT 54.The inner shaft 40 and the outer shaft 50 are concentric and rotateabout the engine axis A. Core airflow is compressed by the LPC 44 thenthe HPC 52, mixed with the fuel and burned in the combustor 56, thenexpanded over the HPT 54 and the ITT 46. The LPT 46 and HPT 54rotationally drive the respective low spool 30 and high spool 32 inresponse to the expansion.

In one non-limiting example, the gas turbine engine 20 is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20bypass ratio is greater than about six (6:1). The geared architecture 48can include an epicyclic gear train, such as a planetary gear system orother gear system. The example epicyclic gear train has a gear reductionratio of greater than about 2.3:1, and in another example is greaterthan about 2.5:1. The geared turbofan enables operation of the low spool30 at higher speeds that can increase the operational efficiency of theLPC 44 and LPT 46 and render increased pressure in a fewer number ofstages.

A pressure ratio associated with the LPT 46 is pressure measured priorto the inlet of the LPT 46 as related to the pressure at the outlet ofthe LPT 46 prior to an exhaust nozzle of the gas turbine engine 20. Inone non-limiting embodiment, the bypass ratio of the gas turbine engine20 is greater than about ten (10:1), the fan diameter is significantlylarger than that of the LPC 44, and the LPT 46 has a pressure ratio thatis greater than about five (5:1). It should be understood, however, thatthe above parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present disclosure is applicable toother gas turbine engines including direct drive turbofans.

In one embodiment, a significant amount of thrust is provided by thebypass flow path B due to the high bypass ratio. The fan section 22 ofthe gas turbine engine 20 is designed for a particular flightcondition—typically cruise at about 0.8 Mach and about 35,000 feet(10,688 meters). This flight condition, with the gas turbine engine 20at its best fuel consumption, is also known as bucket cruise ThrustSpecific Fuel Consumption (TSFC). TSFC is an industry standard parameterof fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of (T/518.7)^(0.5), where “T” represents the ambienttemperature in degrees Rankine. The Low Corrected Fan Tip Speedaccording to one non-limiting embodiment of the example gas turbineengine 20 is less than about 1150 feet per second (351 meters persecond).

Referring to FIG. 2, a single turbine airfoil stage 60 of multiplestages of the HPT 54 is illustrated. The airfoil stage 60 may be a firststage (i.e. leading upstream stage) of the HPT 54 and includes a leadingor upstream, static, vane array 62 and an axially adjacent anddownstream, rotating, blade array (not shown). The vane array 62 has aplurality of circumferentially spaced airfoils 64 (with respect toengine axis A) each extending radially between, ring-shaped, inner andouter endwalls 66, 68 (i.e. platforms). Each vane array 62 may becircumferentially divided into a plurality of vane assemblies 70 eachhaving at least one airfoil 64 and a portion of the inner and outerendwalls 66, 68. When assembled, the vane assemblies 70 forms theannular shape of the vane array 62 concentrically located about theengine axis A.

Referring to FIGS. 2 and 3, the inner endwall 66 has a radially outwardfacing surface 72 that spans axially between fore and aft rims 74, 76 ofthe inner endwall. Similarly, the outer endwall 68 has a radially inwardfacing surface 78 that spans axially between fore and aft rims 80, 82 ofthe outer endwall. The outward surface 72 and the inward surface 78oppose one another and define, in-part, a hot gas flowpath 84 radiallytherebetween for the core airflow C.

Each airfoil 64 of the vane array 62 generally extends through theflowpath 84 to redirect the airflow C that, in-turn is received by thedownstream blade array for converting airflow energy into work generallyrepresented by rotation of the high spool 32. Each airfoil 64 has aconcave pressure side 86 and an opposite convex suction side 88. Thesides 86, 88 span between and generally meet at leading and trailingedges 90, 92 of the airfoil 64.

Each airfoil 64 is circumferentially spaced from the next adjacentairfoil by a pitch distance (see arrow 94) and extends axially by anaxial chord length (see arrow 96). A pitch-to-chord ratio of the firststage vane array 62 (i.e. all vanes of one array) is equal to or greaterthan 1.7, may generally be within a pitch-to-chord ratio range of about1.7 to 2.0, and is preferably about 1.8. With increasing pitch-to-chordratios, flowpath blockage generally decreases. This decrease may beparticularly advantageous for the operating parameters of gearedturbofan engines as described above. In more traditional or conventionalfirst stage vane arrays, pitch-to-chord ratios are lower than 1.7 andmay generally be within a range of 1.20 to 1.68.

The increased pitch-to-chord ratio of the present disclosure, reducesthe number of required airfoils in an array, thus reducing the requiredcooling needs of the HPT 54, which may increase engine operatingefficiency. Furthermore, a decrease in the number of airfoils whencompared to more traditional engines reduces maintenance cost andweight. Alternatively or in addition, the increased pitch-to chord ratiomay generally represent a decrease in chord length 96, which enablesdesigning an HPT with a reduced axial length thereby improving packagingof the entire engine.

Referring to FIG. 3 and as part of a first stage 60 of the HPT 54, eachairfoil 64 of the vane array 62 may have a thickness-to-axial chordratio that is greater than forty percent (40%) and may be aboutfifty-three percent (53%). The “thickness” (see arrow 97) is generallythe maximum thickness of the airfoil, and may be measured generally atabout mid-chord (and/or slightly toward the leading edge 90) of theairfoil 64 and between the sides 86, 88. The “axial chord” is the axialchord length 96 previously described. A higher thickness-to-axial chordratio drives higher Mach numbers and higher losses; however, the higherpitch-to-chord ratio serves to counteract this undesirable effect.

A trailing edge angle (see arrow 99) measured at the trailing edge 92 ofthe airfoil 64 may be greater than about seventy-five (75) degrees.Angle 99 (i.e. trailing edge metal angle) is generally measured betweenan extrapolated line extended from the trailing edge direction and anaxial line generally parallel to the engine axis A. The trailing edgemetal angles of later or aft stages of the HPT 54 are typically lessthan seventy-five (75) degrees.

Referring to FIGS. 3 through 5, higher pitch-to-chord ratios may, insome instances, cause an increase in the formation of air vorticeswithin the airflow C and proximate to the endwalls 66, 68. Such vorticesmay disrupt core airflow C causing a decrease in engine efficiency. Insuch isolated examples, one or both of the outward and inward surfaces72,78 of the respective inner and outer endwalls 66, 68 may carry aplurality of profiled convex and concave regions 98, 100 configured todirect airflow C through the flowpath 84 while minimizing or eliminatingunwanted air vortices.

The convex and concave regions 98, 100 are illustrated in FIG. 5 withtopographic contour lines. Each convex region 98 may be located near andgenerally upstream of a respective leading edge 90 of the airfoils 64.Each concave region 100 may be circumferentially centered between andspaced from opposing pressure and suction sides 86, 88 of adjacentairfoils 64 and substantially centered axially between the leading andtrailing edges 90, 92 (i.e. mid-chord). The convex regions 98 generallyrepresent a projection of the respective surfaces 72, 78 into theflowpath 84 and the concave regions 100 generally represent a depressionof the respective surface 72, 78 that extends away from the flowpath 84(i.e. effectively enlarges the flowpath at the depressions).

Each convex region 98 gradually increases in height to a radial extent102 positioned immediately adjacent to and axially upstream of theleading edge 90 of each airfoil 64. Each concave region 100 graduallyincreases in depth to a radial extent 104. The concave region 100 mayextend axially along a significant percentage of the axial chord length96 (see FIG. 3). The radial extent 104 of the concave region 100 mayextend axially and by an axial distance that is within a range of aboutthirty to eighty percent of the axial chord length 96, and may extendcircumferentially by a circumferential distance that is within a rangeof about thirty to seventy percent of the distance between opposingpressure and suction sides 86, 88 of adjacent airfoils 64 (i.e. pitchdistance 94).

It is further understood and contemplated that the pitch-to-chord ratiotaught in the present disclosure may apply to other stages in theturbine section 28 of the gas turbine engine 20. Furthermore, thepitch-to-chord ratio and/or the convex and concave regions 98, 100taught in the present disclosure may apply to any vane and/or bladearray in any stage of the turbine section 28 or the compressor section24.

While the invention is described with reference to exemplaryembodiments, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted withoutdeparting from the spirit and scope of the invention. In addition,different modifications may be made to adapt the teachings of theinvention to particular situations or materials, without departing fromthe essential scope thereof. The invention is thus not limited to theparticular examples disclosed herein, but includes all embodimentsfalling within the scope of the appended claims.

What is claimed is:
 1. A high pressure turbine of a gas turbine enginecomprising: a first stage vane assembly comprising a first airfoilcircumferentially spaced from an adjacent second airfoil and orientatedabout an engine axis, wherein the first airfoil has a leading edge and atrailing edge with the trailing edge configured to be circumferentiallyseparated from the trailing edge of the second airfoil by a pitchdistance, and the leading and trailing edges of the first airfoil areaxially separated by an axial chord length, and wherein a pitch-to-chordratio of pitch distance over axial chord length is equal to or greaterthan 1.7.
 2. The high pressure turbine set forth in claim 1, wherein thefirst airfoil has a thickness-to-axial chord ratio that is greater thanforty percent.
 3. The high pressure turbine set forth in claim 2,wherein the thickness-to-axial chord ratio is about fifty-three percent.4. The high pressure turbine set forth in claim 1, wherein the trailingedge has an angle that is greater than seventy-five degrees relative tothe engine axis.
 5. The high pressure turbine set forth in claim 4,wherein the pitch-to-chord ratio is within a range of about 1.7 to 2.0.6. The high pressure turbine set forth in claim 1, wherein thepitch-to-chord ratio is within a range of about 1.7 to 2.0.
 7. The highpressure turbine set forth in claim 1, wherein the pitch-to-chord ratiois about 1.8.
 8. The high pressure turbine set forth in claim 1 furthercomprising: an inner endwall, wherein the first airfoil projectsradially outward from an outward surface of the inner endwall and theoutward surface includes at least in-part a concave region locatedaxially between the leading and trailing edges and circumferentiallyadjacent to the first airfoil.
 9. The high pressure turbine set forth inclaim 8, wherein the outward surface includes a convex region proximateto the leading edge of the first airfoil.
 10. The high pressure turbineset forth in claim 1 further comprising: an outer endwall, wherein thefirst airfoil projects radially inward from an inward surface of theouter endwall and the inward surface includes at least in-part a concaveregion located axially between the leading and trailing edges andcircumferentially adjacent to the first airfoil.
 11. The high pressureturbine set forth in claim 10, wherein the inward surface includes aconvex region proximate to the leading edge of the first airfoil.
 12. Agas turbine engine comprising: a low pressure turbine; and a highpressure turbine upstream of the low pressure turbine with respect to acore airflow, the high pressure turbine comprising a vane arrayincluding a plurality of airfoils circumferentially spaced fromone-another and orientated about an engine axis, wherein each one of theplurality of airfoils have a leading edge and a trailing edge with eachone of the trailing edges being circumferentially separated by the nextadjacent trailing edge by a pitch distance, and the leading and trailingedges of each one of the plurality of airfoils are axially separated byan axial chord length, and wherein a pitch-to-chord ratio of pitchdistance over axial chord length is equal to or greater than 1.7.